Blade for a gas turbine

ABSTRACT

Disclosed is a blade for a gas turbine, in particular for an aircraft engine, wherein the blade comprises at least one internal and/or external supporting structure of a metallic and/or intermetallic material. Also disclosed are a blade cluster having at least one such blade for a gas turbine, a method for producing such a blade cluster and a gas turbine.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application claims priority under 35 U.S.C. §119 of German Patent Application No. 102013219774.8, filed Sep. 30, 2013, the entire disclosure of which is expressly incorporated by reference herein.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The invention relates to a blade for a gas turbine. The invention further relates to a blade cluster for a gas turbine, having at least one such blade, a method for producing such a blade cluster and a gas turbine.

2. Discussion of Background Information

Blades are used in various configurations, for example as guide blades or rotor blades, guide grate segments, guide grate rings and the like in gas turbines such as aircraft engines or static industrial gas turbines. Such blades commonly consist of metal alloys or intermetallic compounds. In order to reduce the fuel consumption of gas turbines, a constant improvement in the efficiency of the latter is sought. One possibility for improving efficiency lies for example in raising turbine inlet temperatures. However, this requires the use of suitable high-temperature materials. Therefore, in order to produce blades for gas turbines, high temperature-resistant metallic and intermetallic materials such as nickel- and cobalt-based alloys and the like are used. However, the blades produced in this manner are relatively heavy.

A further possibility for raising the efficiency of gas turbines and accordingly reducing their fuel consumption consists in the use of materials which are as lightweight as possible, such as titanium aluminide compounds which, however, are generally very expensive. It is further known to produce blades not from solid material, but to provide them with cavities. However, because of the given thermal and mechanical loads in the gas turbine, the wall thicknesses cannot be reduced as much as desired.

The present invention has the object of providing a high temperature-resistant and lightweight blade for a gas turbine. Further objects of the invention consist in providing a correspondingly high temperature-resistant and lightweight blade cluster for a gas turbine and a method for producing such a blade cluster and a gas turbine having a high temperature-resistant and lightweight blade or a high temperature-resistant and lightweight blade cluster.

SUMMARY OF THE INVENTION

The objects are achieved according to the invention by a blade, a blade cluster, a method for producing such a blade cluster and by a gas turbine as indicated in the present independent claims. Advantageous configurations with expedient developments of the invention are indicated in the respective dependent claims.

A first aspect of the invention relates to a blade for a gas turbine, in particular for an aircraft engine. In this context, the blade has, according to the invention, an internal and/or external supporting structure made from a metallic and/or intermetallic material. The transfer of forces of the blade then advantageously occurs via the metallic or intermetallic supporting structure.

In one advantageous configuration of the blade according to the invention, the supporting structure is surrounded, at least in certain regions, by a casing made from at least one ceramic fiber composite material. This is in particular the case for that region of the blade which is arranged in the gas duct of the gas turbine during the intended use of the blade, i.e. for the blade airfoil, which typically has a wing profile. In such a configuration of the blade, the transfer of forces also advantageously occurs via the metallic or intermetallic supporting structure, that is to say the core of the blade, which is partially or entirely surrounded with at least one ceramic fiber composite material (Ceramic Matrix Composites, CMC), for example in the form of a shell or part shell. Thus, on one hand, the weight of the blade is reduced, which reduces the fuel consumption of the associated gas turbine. On the other hand, the blade permits higher inlet temperatures and is thus for example suited to high-speed low-pressure turbines. Ceramic fiber composite materials are generally distinguished with respect to the metallic or intermetallic materials mentioned at the outset by having better heat resistance while at the same time being lightweight. For that reason, however, they can often not be subjected to such high mechanical loads, as they are more brittle.

In particular when the supporting structure is fully encased by the ceramic fiber composite material, the supporting structure can also be made of a metallic material which is more lightweight and/or more cost-effective than the metallic or intermetallic material of conventional blades, since, by virtue of the ceramic fiber composite material surrounding it, it is not exposed to such high temperatures. If the metallic or intermetallic material of conventional blades is used for the supporting structure, it is possible to use accordingly less of this material.

Fundamentally, any known CMC material, for example C/C, C/SiC, SiC/SiC or Al₂O₃/Al₂O₃, may be used, individually or in any combination, as the ceramic fiber composite material. The blade according to the invention may, by virtue of its lower weight, be designed for higher structural-mechanical limits, for example increased oscillating loading, while at the same time subjecting inner and outer shrouds, hubs, suspensions and the like to lower structural-mechanical load. This results in additional advantages for adjoining components and component groups of an associated gas turbine, since these can also be accordingly more lightweight.

It is to be noted in general that the term “casing” in the sense of the present invention is not to be understood as a simple coating of a metallic or intermetallic blade or of a blade airfoil. Such coatings, as known from the prior art, are very thin, i.e. they are generally only a few micrometers thick. In contrast, according to the present invention, “casing” is to be understood as a substantially thicker shell or shell section, in particular at least half a millimeter thick, preferably at least one millimeter or several millimeters thick.

In order to produce the blade according to the invention, an internal and/or external supporting structure, which is designed to be skeletal or which consists of a grid-like and/or bionic structure, can first be forged. In other words, the erstwhile blade airfoil is reduced to a functional structural-mechanical supporting element. This can, fundamentally, adopt any suitable geometry or bionic structure, since it has only to provide a type of backbone (or backbone brace) and/or outer supporting structure in order to take up, elastically, the forces arising when the gas turbine is in operation. Attaching a CMC shell, a CMC part shell or a CMC blade element to and/or within the supporting structure completes the blade. To that end, for example, a CMC shell as a casing can be slotted over the supporting structure serving as a backbone and can be connected thereto via an outward- and/or inward-oriented form-fitting plug connection. In addition or as an alternative, the supporting structure and the casing can preferably also be connected in a material-bonded manner by soldering, welding and/or adhesive bonding. As an alternative or in addition, it can also be provided that the CMC material is cast onto or into the supporting structure, thus achieving a mechanically particularly stable bond between the casing and the supporting structure. Then, the blade can, where necessary, be provided with a shroud. Fundamentally, the casing can partially or entirely sheathe the supporting structure.

In one advantageous embodiment of the invention, the supporting structure is, at least in certain regions, solid and/or pierced and/or grid-like and/or porous. This permits in particular the weight and the mechanical properties of the blade according to the invention to be optimally matched to its respective intended use.

A further advantageous embodiment of the invention provides that the supporting structure bears, at least in certain regions, directly on an inner side of the casing and/or wherein at least one cavity is arranged between the supporting structure and the casing. With the supporting structure bearing directly on the inner side of the casing, it is possible to achieve correspondingly good support and particularly high mechanical stability. Conversely, it is possible, by providing one or more cavities between the supporting structure and the casing, to further reduce the weight of the blade. The cavity or cavities may additionally be acted upon with coolant so as to internally cool the blade.

Further advantages result when the supporting structure is supported by at least one supporting element on an inner wall of the casing. With the aid of a supporting element, the casing can be supported in a targeted manner at regions which are subjected to particularly high mechanical load. It can furthermore be provided that the casing has a groove or recess which corresponds to the at least one supporting element. As an alternative or in addition, it is provided that the casing is slotted onto the supporting structure. This permits particularly simple production of the blade or of the blade profile in the region of the blade airfoil. Any supporting elements which are present may then advantageously serve as positioning or insertion aids, by cooperation with any grooves which are present in the casing.

A further advantageous embodiment of the invention provides that the casing forms a substantially convex suction side and/or a substantially concave pressure side and/or a leading edge connecting the suction side and the pressure side and/or a trailing edge connecting the suction side and the pressure side of a blade airfoil of the blade. In other words, it is provided that the casing forms one, several or preferably all of the aerodynamic surfaces of the blade or of the blade airfoil. This also allows the blade to be optimally matched to various intended uses and to the mechanical and thermal loads on the blade which arise during operation.

Further advantages result when the casing comprises at least one end region which projects away from the supporting structure so as to cover, at least in certain regions, a component of the gas turbine and/or so as to bind to an adjoining casing of a further blade and/or so as to line, at least in certain regions, a flow duct of the gas turbine. By virtue of such a projecting end region, the casing of the supporting structure can advantageously be used to cover or line further components of the gas turbine.

A further advantageous embodiment of the invention provides that the blade is connected to a hub and/or is formed as a rotor blade and/or as a guide blade and/or is present as a blade cluster together with at least one further blade. The blade can thus be used in a particularly variable manner for various purposes in various gas turbines. Furthermore, the blade can be used as a guide blade for producing a guide grate or a guide grate segment or a guide grate ring. Equally, the blade can be formed as a rotor blade of a rotor. Furthermore, the blade can be present together with at least one further blade as a blade cluster. The at least one further blade can then be formed in a fundamentally conventional manner or also according to the invention.

A rotor blade typically has a blade root, which may for example have a dovetail or fir-tree shape, in order to be inserted into and held in a groove of complementary form in a disk which is connected to a rotor shaft. Radially outward, an inner shroud or inner platform of the blade generally adjoins the blade root. Further radially outward, the actual blade airfoil with an aerodynamic profile adjoins the inner shroud or platform, which airfoil has, according to the invention, a metallic or intermetallic supporting structure and, at least partially surrounding this supporting structure, a shell or part shell made of a ceramic fiber composite material. Further radially outward, the blade may also have an outer shroud or platform.

A guide blade has, in contrast with a rotor blade, no blade root, but may be provided, in a similar manner to the rotor blade, with an inner shroud or platform and/or an outer shroud or platform. In addition, the guide blade also has a blade airfoil, which can be constructed as described above.

In terms of production technology, the shrouds may be formed separately from the corresponding blade airfoil of the guide blade or rotor blade and may, if appropriate, be coated for protection from erosion and/or oxidation. The shrouds can then be connected in a force-fitting and/or form-fitting manner to the corresponding blade airfoil. The shrouds can be produced either from a metallic or intermetallic material or from a ceramic material, in particular from a ceramic fiber composite material. The shrouds and/or the blade airfoil can be produced by generative or additive methods and/or by wrapping and/or cross-laying in molds. Special looms are nowadays able to weave ceramic fibers, in a similar manner to textile fibers. It is also conceivable to produce the entire blade almost in one piece by means of generative or additive methods, wherein in this case, in certain regions, metallic or intermetallic and ceramic materials are preferably to be melted using a laser.

A second aspect of the invention relates to a blade cluster for a gas turbine, in particular for an aircraft engine. Here the blade cluster comprises at least two blades connected to one another. At least one of the blades may then be formed according to the first invention aspect. As an alternative or in addition, it is also possible for at least one of the blades of the blade cluster to be formed at least partially from at least one ceramic composite material, whereas at least one other blade of the blade cluster may be formed from a metallic and/or intermetallic material and may be substantially free from ceramic composite material. “Substantially free from ceramic composite material” means in this context that the blades or blade airfoils made of a metallic or intermetallic material may perfectly well be coated with a ceramic composite material, wherein a layer of such a coating is generally, as already mentioned above, substantially less than one millimeter thick.

The features resulting from this—and the advantages thereof—may be taken from the descriptions of the first invention aspect, wherein advantageous configurations of the first invention aspect are to be regarded as advantageous configurations of the second invention aspect and vice versa. Fundamentally, some or all of the blades of the blade cluster may be formed according to the first invention aspect. In the event that in the blade cluster one or more blades, at least in the region of their blade airfoil, is/are formed exclusively from ceramic composite material, whereas all the remaining blades, at least in the region of their blade airfoil, are formed exclusively from a metallic or intermetallic material, these remaining blades or blade airfoils take on the function according to the invention of the supporting structure for the ceramic blades or blade airfoils of the blade cluster.

In one advantageous embodiment of the invention, the blades of the blade cluster are connected to an inner platform and/or an outer platform. It is thus possible for the blade cluster to be optimally adapted, with respect to its mechanical properties and with respect to its flow properties, to its respective intended use. The inner and/or outer platform is then preferably connected at least to the supporting structure of the blade according to the invention or of the blade airfoil. The inner and/or the outer platform can then at the same time function as a common attachment means of the blades of the blade cluster or be itself connected to such an attachment means.

Further advantages emerge when the blade cluster comprises an odd number of blades and/or in alternation one blade which consists entirely of a metallic and/or intermetallic material, and one blade formed according to the first invention aspect. Where the blade cluster has, in alternation, a conventional blade and a blade according to the invention, a hybrid blade cluster is obtained which enables a low-risk, systematic approximation to a blade cluster made entirely from CMC. It is also possible to use a blade or a blade airfoil which is formed exclusively from a ceramic fiber composite material.

A further advantageous configuration of the invention provides that mutually adjoining casings of at least two blades are connected to each other in a form-fitting and/or material-bonded manner As a result, it is possible to obtain, between the two blades or blade airfoils, a surface which can be subjected to thermal and mechanical load.

Further advantages emerge when the blade cluster comprises a covering of the inner and/or outer shroud or platform and/or a housing covering and/or a lining for a flow duct made from at least one ceramic fiber composite material. This also presents a constructively advantageous possibility of preparing surfaces which can be subjected to high thermal and mechanical load.

For that purpose, a further embodiment has proven to be advantageous if the already present CMC casing of the at least one blade according to the invention is formed such that it can also be used for example as a shroud covering, housing covering and/or flow duct covering or lining.

A third aspect of the invention relates to a method for producing a blade cluster for a gas turbine, in particular for an aircraft engine. In that context, it is provided according to the invention that at least one metallic and/or intermetallic internal and/or external supporting structure of at least one blade is connected to a, preferably metallic and/or intermetallic, inner platform and to a, preferably metallic and/or intermetallic, outer platform, and/or wherein at least one of the blades of the blade cluster is at least partially formed from at least one ceramic composite material, whereas at least one other blade of the blade cluster is formed from a metallic and/or intermetallic material and is substantially free from ceramic composite material. It can thus be provided that at least one blade of the blade cluster is produced such that it has, in the sense of the first and second invention aspects, a supporting structure made from a metallic and/or intermetallic material.

The supporting structure can in turn, at least in certain regions, be sheathed with a casing made from at least one ceramic fiber composite material. The supporting structure consisting of a metallic and/or intermetallic material is connected to an also metallic and/or intermetallic inner and outer platform and can in turn, at least in certain regions, be sheathed with at least one CMC material. A high temperature-resistant and lightweight blade cluster having at least one hybrid CMC blade is thus obtained. In that context, it can fundamentally be provided that, in addition to this hybrid CMC blade, one or more conventionally produced blades are provided and are connected to the inner and outer platform. Alternatively, some or all of the blades may be formed as hybrid CMC blades. Further features and the advantages resulting therefrom may be taken from the descriptions of the first and second invention aspects, wherein advantageous configurations of the first and second invention aspects are to be regarded as advantageous configurations of the third invention aspect and vice versa.

One advantageous embodiment of the invention provides that the inner platform and at least one supporting structure are first produced in a composite assembly, after which the casing is slotted onto the supporting structure and the supporting structure is connected to the outer platform.

Alternatively or in addition, it is provided that the outer platform and at least one supporting structure are first produced in a composite assembly, after which the casing is slotted onto the supporting structure and the supporting structure is connected to the inner platform. Alternatively or in addition, it is provided that the inner platform, the outer platform and at least one supporting structure are first produced independently from one another, are connected to one another and the supporting structure is then provided with the casing. It is hereby possible to produce the at least one hybrid blade and thus the blade cluster in a particularly flexible fashion. In that context, it can fundamentally be provided that various blades are produced identically or differently.

Further advantages emerge when the casing of the at least one blade is connected to the inner platform and/or to the outer platform by means of a plug connection and/or a clip connection. The casing can thus be assembled and disassembled in a particularly quick, simple and cost-effective manner

A fourth aspect of the invention relates to a gas turbine, in particular an aircraft engine, which comprises at least one blade formed in accordance with the first invention aspect and/or at least one blade cluster which is formed in accordance with the second invention aspect, and/or is produced by means of a method in accordance with the third invention aspect. This provides a gas turbine in which the transfer of forces of the blades occurs advantageously via the metallic or intermetallic supporting structure(s). The features resulting from this—and the advantages thereof—may be taken from the descriptions of the first, second and third invention aspects, wherein advantageous configurations of the first, second and third invention aspects are to be regarded as advantageous configurations of the fourth invention aspect and vice versa.

BRIEF DESCRIPTION OF THE DRAWINGS

Further features of the invention emerge from the claims, the exemplary embodiments and with reference to the drawings. The features and combinations of features mentioned above in the description, and the features and combinations of features mentioned below in the exemplary embodiments, may be employed not only in the combination indicated in each case, but may also be employed in other combinations without departing from the scope of the invention. In the figures:

FIG. 1 shows a schematic perspective view of a blade cluster according to the invention;

FIG. 2 shows a schematic cross section through a blade according to the invention;

FIG. 3 shows a schematic longitudinal section through a hub having multiple blades according to the invention;

FIG. 4 shows a schematic longitudinal section through a hub, to which is attached an embodiment of the blade according to the invention;

FIG. 5 shows a schematic longitudinal section through a hub, to which is attached a further embodiment of the blade according to the invention;

FIG. 6 shows a schematic longitudinal section through a hub, to which is attached a further embodiment of the blade according to the invention; and

FIG. 7 shows a schematic perspective view of a blade cluster according to the invention, according to a further exemplary embodiment.

DETAILED DESCRIPTION OF THE PRESENT INVENTION

The particulars shown herein are by way of example and for purposes of illustrative discussion of the embodiments of the present invention only and are presented in the cause of providing what is believed to be the most useful and readily understood description of the principles and conceptual aspects of the present invention. In this regard, no attempt is made to show details of the present invention in more detail than is necessary for the fundamental understanding of the present invention, the description in combination with the drawings making apparent to those of skill in the art how the several forms of the present invention may be embodied in practice.

FIG. 1 shows a schematic perspective view of a blade cluster 10, according to the invention, for an aircraft engine. In this exemplary embodiment, a guide blade cluster 10 is shown. Alternatively, however, the invention could be demonstrated just as well for example with reference to a rotor blade cluster. The blade cluster 10 shown comprises five blades which are exemplary in number and in arrangement and of which the outer blades 12 and the central blade 12 have in each case a supporting structure 14 made of a metallic and/or intermetallic material. These supporting structures 14 are in each case surrounded by a casing 16 made of at least one ceramic fiber composite material (CMC). For the sake of clarity, the casing 16 of the central blade 12 is only partially represented, such that the supporting structure 14 is free in certain regions. Furthermore, two conventional blades 18, which each consist entirely of a metallic and/or intermetallic material, are arranged between the blades 12. For reasons of clarity, these blades 18 are also only partially represented.

The supporting structures 14 are in each case connected to an inner platform 20 and an outer platform 22 of the blade cluster 10, in order to transfer and distribute forces and in order to damp vibrations when the associated aircraft engine is in operation. By contrast, the CMC casings 16 are inserted into corresponding recesses 32 (FIG. 4) in the inner platform 20 and the outer platform 22. In addition to the shown 1,3,5 arrangement of the hybrid CMC blades 12, other arrangements, for example 1,5, are also possible. In the present case, the CMC blades 12 thus replace three conventional standard blades 18 or blade airfoils. The two remaining standard blades 18 are retained in order to maintain the structural-mechanical integrity of the blade cluster 10. Moreover, the blade cluster 10 may be combined with identically or differently constructed blade clusters 10 and/or conventional blade clusters (not shown), that is to say with blade clusters which exclusively comprise blades 18 made of metal and/or an intermetallic material, so as to form a blade ring.

In order to produce the blade cluster 10, first the inner platform 20, the supporting structures 14, the conventional blades 18 and the outer platform 22 are prepared. In this context, either the inner platform 20 and the supporting structures 14 or the conventional blades 18 can be produced already in a composite assembly and the outer platform 22 as a separate component. Alternatively, the outer platform 22 and the supporting structures 14 or the conventional blades 18 can be provided in a composite assembly and the inner platform 20 as a separate component. Equally, all of the abovementioned components may be prepared separately. The supporting structures 14, that is to say the erstwhile blade airfoils, are, in comparison with the conventional blades 18, reduced to functional structural-mechanical elements. The supporting structures 14 may then fundamentally have any suitable geometry or bionic structure, since they only represent a form of backbone (or backbone brace) and in later operation take up the forces. The casings 16, which function as aerodynamic elements, may be made as shells made of CMC and be slotted over the backbone of the supporting structures 14. Depending on the configuration of the abovementioned components, the blades 12 and 18 are connected outwardly and/or inwardly to the inner platform 20 or the outer platform 22 (shroud) via an outwardly or inwardly oriented form-fitting plug connection. The connection may in addition be executed in a force-fitting or material-bonded manner by means of soldering, welding or adhesive bonding. A connection of this type is in particular also advantageous for aerodynamic reasons, in order to close gaps.

A single blade 12, with or without a shroud or outer platform 22, may be produced in a similar manner For example, first a blade pedestal (not shown) having a skeletal or grid-like (bionic structure) supporting structure 14 may be forged, cast and/or produced generatively. By the introduction or production of a casing 16 which serves as a blade airfoil, and where appropriate by connecting the casing 16 to the shroud, the blade 12 is completed in a manner similar to that described above.

FIG. 2 shows a schematic cross section through a blade 12 according to the invention, along the plane of section II shown in FIG. 1. It is clear that the supporting structure 14 is supported on an internal wall of the casing 16 via three integral supporting elements 24 which are exemplary in terms of number and arrangement. Thus, the supporting structure 14 does not entirely fill the internal space of the CMC casing 16, such that weight-saving cavities 26 remain, through which for example cooling air can be guided. Alternatively, it may also be provided that the supporting structure 14 entirely fills the casing 16. The supporting elements 24 contribute, inter alia, to the prevention or at least the reduction of vibrations in the CMC casing 16.

FIG. 3 shows a schematic longitudinal section through a metallic hub 28 (or the inner platform 20), to which are attached multiple blades 12 according to the invention. In contrast to the embodiment mentioned above, according to FIG. 4, it will be noted that in the case of this embodiment the CMC casings 16 of the blades 12 have in each case an end region 30 which projects away from the supporting structure 14. Mutually adjoining end regions 30 are secured to one another for example in a material-bonded manner or via clip connections. Thus, the flow duct of the aircraft turbine, which lies between the blades 12, is lined with the CMC casings 16 such that it is protected, at least in certain regions, from hot gases; cover the hub 28 and said CMC casings 16 also simplify the binding of mutually adjoining casings 16.

FIG. 4 to FIG. 6 respectively show schematic longitudinal sections through hubs 28 (or inner platforms 20), to which are respectively attached, in various manners, various embodiments of the blade 12 according to the invention. In the embodiment shown in FIG. 4, the CMC casings 16 of the blade 12 are arranged in a corresponding recess 32 in the hub 28. In this embodiment, no supporting structure 14 is represented. This can either be imagined or this blade can be a blade whose blade airfoil is exclusively formed from CMC. In the latter case, at least one further blade of the blade cluster, whose blade airfoil is conventionally formed from a metallic or intermetallic material, then takes on the function of the supporting structure.

In FIG. 5, the hub 28 (or inner platform 20) comprises a connection region 34, onto which the CMC casings 16 are slotted on one side and which on the other side is connected in a material-bonded manner to the supporting structure 14. The casing 16 bears against the hub 28 (or the inner platform 20).

Finally, FIG. 6 shows an embodiment in which the casing 16 is slotted onto the supporting structure 14 and bears against the connection region 34 of the hub 28 (or inner platform 20). However, many other binding variants may also be provided, which may be applied individually or in any combination.

FIG. 7 shows a schematic perspective view of a blade cluster 10 according to the invention, according to a further exemplary embodiment. In contrast with the preceding exemplary embodiment, only two of the total five blades according to the invention are formed as blades 12 having a CMC blade airfoil, while the remaining blades are formed as conventional blades 18. The outer blade 12, which is depicted in a partially transparent manner, comprises a grid-like supporting structure 14 in the region of its blade airfoil, while the central blade 12 has, in the region of its blade airfoil, a CMC casing 16 which is entirely filled by the supporting structure 14. Alternatively, the central blade 12 could, in the region of its blade airfoil, also be formed exclusively by the CMC casing 16, i.e. have no metallic or intermetallic supporting structure connecting the inner platform 20 to the outer platform 22. In this case, at least one of the other blades of the blade cluster then take on the function of the supporting structure.

While the present invention has been described with reference to exemplary embodiments, it is understood that the words which have been used herein are words of description and illustration, rather than words of limitation. Changes may be made, within the purview of the appended claims, as presently stated and as amended, without departing from the scope and spirit of the present invention in its aspects. Although the present invention has been described herein with reference to particular means, materials and embodiments, the present invention is not intended to be limited to the particulars disclosed herein; rather, the present invention extends to all functionally equivalent structures, methods and uses, such as are within the scope of the appended claims. 

1-19. (canceled)
 20. A blade for a gas turbine, wherein the blade comprises at least one internal and/or external supporting structure of a metallic and/or intermetallic material.
 21. The blade of claim 20, wherein the supporting structure is surrounded, at least in certain regions, by a casing of at least one ceramic fiber composite material.
 22. The blade of claim 20, wherein the supporting structure is, at least in certain regions, solid and/or pierced and/or grid-like and/or porous.
 23. The blade of claim 21, wherein the supporting structure bears, at least in certain regions, directly on an inner side of the casing and/or at least one cavity is arranged between the supporting structure and the casing.
 24. The blade of claim 21, wherein the supporting structure is supported by at least one supporting element on an inner wall of the casing and/or the casing is slotted onto the supporting structure.
 25. The blade of claim 21, wherein the casing forms a substantially convex suction side and/or a substantially concave pressure side and/or a leading edge connecting the suction side and the pressure side and/or a trailing edge connecting the suction side and the pressure side of a blade airfoil of the blade.
 26. The blade of claim 21, wherein the casing comprises at least one end region which projects away from the supporting structure so as to cover, at least in certain regions, a component of the gas turbine and/or so as to bind to an adjoining casing of a further blade and/or so as to line, at least in certain regions, a flow duct of the gas turbine.
 27. The blade of claim 20, wherein the blade is connected to a hub and/or is formed as a rotor blade and/or as a guide blade and/or is present as a blade cluster together with at least one further blade.
 28. A blade cluster for a gas turbine, wherein the blade cluster comprises at least two blades connected to one another, at least one of the blades being the blade of claim 20, and/or at least one of the blades of the blade cluster being formed at least partially from at least one ceramic composite material and at least one other blade of the blade cluster is formed from a metallic and/or intermetallic material and is substantially free from ceramic composite material.
 29. The blade cluster of claim 28, wherein the at least two blades are connected to an inner platform and/or to an outer platform.
 30. The blade cluster of claim 28, wherein the blade cluster comprises an odd number of blades and/or comprises in alternation one blade which consists entirely of a metallic and/or intermetallic material and one blade that comprises at least one internal and/or external supporting structure of a metallic and/or intermetallic material.
 31. The blade cluster of claim 28, wherein the supporting structure of a blade is surrounded, at least in certain regions, by a casing of at least one ceramic fiber composite material and mutually adjoining casings of at least two blades are connected to each other in a form-fitting and/or material-bonded manner.
 32. The blade cluster of claim 28, wherein the cluster further comprises a covering of the inner and/or outer shroud or platform and/or a housing covering and/or a lining for a flow duct of at least one ceramic fiber composite material.
 33. The blade cluster of claim 32, wherein a hub covering and/or the housing covering and/or the lining of the flow duct is formed, at least in certain regions, by a casing of at least one blade.
 34. A method for producing a blade cluster for a gas turbine in which at least two blades are connected to one another, wherein the method comprises connecting at least one metallic and/or intermetallic internal and/or external supporting structure of at least one blade to an inner platform of the blade cluster and to an outer platform of the blade cluster and/or comprises at least partially forming at least one of the blades of the blade cluster from at least one ceramic composite material and forming at least one other blade of the blade cluster from a metallic and/or intermetallic material that is substantially free from ceramic composite material.
 35. The method of claim 34, wherein the supporting structure is, at least in certain regions, sheathed with a casing made from at least one ceramic fiber composite material.
 36. The method of claim 35, wherein the inner platform and at least one supporting structure are first produced in a composite assembly, after which the casing is slotted onto the supporting structure and the supporting structure is connected to the outer platform; and/or the outer platform and at least one supporting structure are first produced in a composite assembly, after which the casing is slotted onto the supporting structure and the supporting structure is connected to the inner platform; and/or the inner platform, the outer platform and at least one supporting structure are first produced independently from one another, are connected to one another and the supporting structure is subsequently provided with the casing.
 37. The method of claim 35, wherein the casing of the at least one blade is connected to the inner platform and/or to the outer platform via a plug connection and/or a clip connection.
 38. A gas turbine, wherein the gas turbine comprises the blade of claim
 20. 39. A gas turbine, wherein the gas turbine comprises the blade cluster of claim
 28. 